US20240076050A1

US20240076050A1 - Engine mount system for a gas turbine engine

Info

Publication number: US20240076050A1
Application number: US18/215,394
Authority: US
Inventors: Gabriel L. Suciu; Brian D. Merry; Christopher M. Dye; Steven B. Johnson; Frederick M. Schwarz
Original assignee: RTX Corp
Current assignee: RTX Corp
Priority date: 2008-06-02
Filing date: 2023-06-28
Publication date: 2024-03-07
Also published as: US11286883B2; US20150300207A1; US20150315977A1; US20210372349A1; US20160032828A1; US20170058830A1; US11731773B2; US20140174056A1; US20200284222A1; US10451004B2

Abstract

A gas turbine engine includes, among other things, a propulsor section including a rotor and blades, a gear train, a compressor section and a turbine section. A static structure includes a first case and a second case. The first case includes a first engine mount location. The second case includes a second engine mount location. The first engine mount location is axially near the gear train.

Description

CROSS-REFERENCE TO RELATED APPLICATION

The present disclosure is a continuation of U.S. patent application Ser. No. 17/395,553, filed Aug. 6, 2021, which is a continuation of U.S. patent application Ser. No. 16/405,149, filed May 7, 2019, which is a continuation of U.S. patent application Ser. No. 15/173,288, filed Jun. 3, 2016, which is a continuation of U.S. patent application Ser. No. 14/755,221, filed Jun. 30, 2015, which is a continuation of U.S. patent application Ser. No. 14/190,429, filed Feb. 26, 2014, which was a continuation-in-part of U.S. patent application Ser. No. 13/340,988, filed Dec. 30, 2011, which was a continuation-in-part of U.S. patent application Ser. No. 12/131,876, filed Jun. 2, 2008.

BACKGROUND

The present invention relates to a gas turbine engine and more particularly to an engine mounting configuration for the mounting of a turbofan gas turbine engine to an aircraft pylon.
A gas turbine engine may be mounted at various points on an aircraft such as a pylon integrated with an aircraft structure. An engine mounting configuration ensures the transmission of loads between the engine and the aircraft structure. The loads typically include the weight of the engine, thrust, aerodynamic side loads, and rotary torque about the engine axis. The engine mount configuration must also absorb the deformations to which the engine is subjected during different flight phases and the dimensional variations due to thermal expansion and retraction.
One conventional engine mounting configuration includes a pylon having a forward mount and an aft mount with relatively long thrust links which extend forward from the aft mount to the engine intermediate case structure. Although effective, one disadvantage of this conventional type mounting arrangement is the relatively large “punch loads” into the engine cases from the thrust links which react the thrust from the engine and couple the thrust to the pylon. These loads tend to distort the intermediate case and the low pressure compressor (LPC) cases. The distortion may cause the clearances between the static cases and rotating blade tips to increase which may negatively affect engine performance and increase fuel burn.

SUMMARY

A gas turbine engine according to an exemplary aspect of the present disclosure includes a core nacelle defined about an engine centerline axis, a fan nacelle mounted at least partially around the core nacelle to define a fan bypass airflow path for a fan bypass airflow, a gear train defined along an engine centerline axis, the gear train defines a gear reduction ratio of greater than or equal to about 2.3, a spool along the engine centerline axis which drives the gear train, the spool includes a three to six (3-6) low pressure turbine, and a fan variable area nozzle axially movable relative to the fan nacelle to vary a fan nozzle exit area and adjust a pressure ratio of the fan bypass airflow during engine operation.
In a further non-limiting embodiment of any of the foregoing gas turbine engine embodiments, the gear train may define a gear reduction ratio of greater than or equal to about 2.5.
In a further non-limiting embodiment of any of the foregoing gas turbine engine embodiments, the engine may further include a controller operable to control the fan variable area nozzle to vary the fan nozzle exit area and adjust the pressure ratio of the fan bypass airflow.
In a further non-limiting embodiment of any of the foregoing gas turbine engine embodiments, the controller may be operable to reduce the fan nozzle exit area at a cruise flight condition. Additionally or alternatively, the controller may be operable to control the fan nozzle exit area to reduce a fan instability.
In a further non-limiting embodiment of any of the foregoing gas turbine engine embodiments, the fan variable area nozzle may define a trailing edge of the fan nacelle.
In a further non-limiting embodiment of any of the foregoing gas turbine engine embodiments, the gear train may drive a fan within the fan nacelle.
In a further non-limiting embodiment of any of the foregoing gas turbine engine embodiments, the low pressure turbine may be a five (5) stage low pressure turbine.
In a featured embodiment, a gas turbine engine comprises a gear train defined along an axis. A spool along the axis drives the gear train and includes a low stage count low pressure turbine. A fan is rotatable at a fan speed about the axis and driven by the low pressure turbine through the gear train. The fan speed is less than a speed of the low pressure turbine. A core is surrounded by a core nacelle defined about the axis. A fan nacelle is mounted at least partially around the core nacelle to define a fan bypass airflow path for a fan bypass airflow. A bypass ratio defined by the fan bypass passage airflow divided by airflow through the core is greater than about ten (10).
In another embodiment according to the previous embodiment, the low stage count includes six or fewer stages.
In another embodiment according to any of the previous embodiments, the low pressure turbine is one of three turbine rotors. The low pressure turbine drives the fan, while the other two of the turbine rotors each drive a compressor section.
In another embodiment according to any of the previous embodiments, a high pressure turbine is also included, with each of the low pressure turbine and the high pressure turbine driving a compressor rotor.
In another embodiment according to any of the previous embodiments, the gear train is positioned intermediate a compressor rotor driven by the low pressure turbine and the fan.
In another embodiment according to any of the previous embodiments, the gear train is positioned intermediate the low pressure turbine and the compressor rotor is driven by the low pressure turbine.
Although the different examples have the specific components shown in the illustrations, embodiments of this invention are not limited to those particular combinations. It is possible to use some of the components or features from one of the examples in combination with features or components from another one of the examples.
These and other features disclosed herein can be best understood from the following specification and drawings, the following of which is a brief description.

BRIEF DESCRIPTION OF THE DRAWINGS

The various features and advantages of this invention will become apparent to those skilled in the art from the following detailed description of the currently disclosed embodiment. The drawings that accompany the detailed description can be briefly described as follows:

FIG. 1A is a general schematic sectional view through a gas turbine engine along the engine longitudinal axis;

FIG. 1B is a general sectional view through a gas turbine engine along the engine longitudinal axis illustrating an engine static structure case arrangement on the lower half thereof;

FIG. 1C is a side view of an mount system illustrating a rear mount attached through an engine thrust case to a mid-turbine frame between a first and second bearing supported thereby;

FIG. 1D is a forward perspective view of an mount system illustrating a rear mount attached through an engine thrust case to a mid-turbine frame between a first and second bearing supported thereby;

FIG. 2A is a top view of an engine mount system;

FIG. 2B is a side view of an engine mount system within a nacelle system;

FIG. 2C is a forward perspective view of an engine mount system within a nacelle system;

FIG. 3 is a side view of an engine mount system within another front mount;

FIG. 4A is an aft perspective view of an aft mount;

FIG. 4B is an aft view of an aft mount of FIG. 4A;

FIG. 4C is a front view of the aft mount of FIG. 4A;

FIG. 4D is a side view of the aft mount of FIG. 4A;

FIG. 4E is a top view of the aft mount of FIG. 4A;

FIG. 5A is a side view of the aft mount of FIG. 4A in a first slide position; and

FIG. 5B is a side view of the aft mount of FIG. 4A in a second slide position.

FIG. 6 shows another embodiment.

FIG. 7 shows yet another embodiment.

DETAILED DESCRIPTION

FIG. 1A illustrates a general partial fragmentary schematic view of a gas turbofan engine 10 suspended from an engine pylon 12 within an engine nacelle assembly N as is typical of an aircraft designed for subsonic operation.
The turbofan engine 10 includes a core engine within a core nacelle C that houses a low spool 14 and high spool 24. The low spool 14 includes a low pressure compressor 16 and low pressure turbine 18. The low spool 14 drives a fan section 20 connected to the low spool 14 either directly or through a gear train 25.
The high spool 24 includes a high pressure compressor 26 and high pressure turbine 28. A combustor 30 is arranged between the high pressure compressor 26 and high pressure turbine 28. The low and high spools 14, 24 rotate about an engine axis of rotation A.
The engine 10 in one non-limiting embodiment is a high-bypass geared architecture aircraft engine. In one disclosed, non-limiting embodiment, the engine 10 bypass ratio is greater than about six (6), with an example embodiment being greater than about ten (10), the gear train 25 is an epicyclic gear train such as a planetary gear system or other gear system with a gear reduction ratio of greater than about 2.3 and the low pressure turbine 18 has a pressure ratio that is greater than about 5. In one disclosed embodiment, the engine 10 bypass ratio is greater than ten (10:1), the turbofan diameter is significantly larger than that of the low pressure compressor 16, and the low pressure turbine 18 has a pressure ratio that is greater than 5:1. The gear train 25 may be an epicycle gear train such as a planetary gear system or other gear system with a gear reduction ratio of greater than about 2.5:1. It should be understood, however, that the above parameters are only exemplary of one embodiment of a geared architecture engine and that the present invention is applicable to other gas turbine engines including direct drive turbofans.
Airflow enters the fan nacelle F which at least partially surrounds the core nacelle C. The fan section 20 communicates airflow into the core nacelle C to the low pressure compressor 16. Core airflow compressed by the low pressure compressor 16 and the high pressure compressor 26 is mixed with the fuel in the combustor 30 where is ignited, and burned. The resultant high pressure combustor products are expanded through the high pressure turbine 28 and low pressure turbine 18. The turbines 28, 18 are rotationally coupled to the compressors 26, 16 respectively to drive the compressors 26, 16 in response to the expansion of the combustor product. The low pressure turbine 18 also drives the fan section 20 through gear train 25. A core engine exhaust E exits the core nacelle C through a core nozzle 43 defined between the core nacelle C and a tail cone 33.
With reference to FIG. 1B, the low pressure turbine 18 includes a low number of stages, which, in the illustrated non-limiting embodiment, includes three turbine stages, 18A, 18B, 18C. The gear train 25 operationally effectuates the significantly reduced number of stages within the low pressure turbine 18. The three turbine stages, 18A, 18B, 18C facilitate a lightweight and operationally efficient engine architecture. It should be appreciated that a low number of stages contemplates, for example, three to six (3-6) stages. Low pressure turbine 18 pressure ratio is pressure measured prior to inlet of low pressure turbine 18 as related to the pressure at the outlet of the low pressure turbine 18 prior to exhaust nozzle.
Thrust is a function of density, velocity, and area. One or more of these parameters can be manipulated to vary the amount and direction of thrust provided by the bypass flow B. The Variable Area Fan Nozzle (“VAFN”) 42 operates to effectively vary the area of the fan nozzle exit area 45 to selectively adjust the pressure ratio of the bypass flow B in response to a controller (not shown). Low pressure ratio turbofans are desirable for their high propulsive efficiency. However, low pressure ratio fans may be inherently susceptible to fan stability/flutter problems at low power and low flight speeds. The VAFN 42 allows the engine to change to a more favorable fan operating line at low power, avoiding the instability region, and still provide the relatively smaller nozzle area necessary to obtain a high-efficiency fan operating line at cruise.
A significant amount of thrust is provided by the bypass flow B due to the high bypass ratio. The fan section 20 of the engine 10 is designed for a particular flight condition—typically cruise at about 0.8 Mach and about 35,000 feet. The flight condition of 0.8 Mach and 35,000 ft, with the engine at its best fuel consumption—also known as “bucket cruise Thrust Specific Fuel Consumption (‘TSFC’)”—is the industry standard parameter of lbm of fuel being burned divided by lbf of thrust the engine produces at that minimum point. “Low fan pressure ratio” is the pressure ratio across the fan blade alone, without the Fan Exit Guide Vane (“FEGV”) system 36. The low fan pressure ratio as disclosed herein according to one non-limiting embodiment is less than about 1.45. “Low corrected fan tip speed” is the actual fan tip speed in ft/sec divided by an industry standard temperature correction of [(Tambient deg R)/518.7){circumflex over ( )}0.5]. The “Low corrected fan tip speed” as disclosed herein according to one non-limiting embodiment is less than about 1150 ft/second.
As the fan blades within the fan section 20 are efficiently designed at a particular fixed stagger angle for an efficient cruise condition, the VAFN 42 is operated to effectively vary the fan nozzle exit area 45 to adjust fan bypass air flow such that the angle of attack or incidence on the fan blades is maintained close to the design incidence for efficient engine operation at other flight conditions, such as landing and takeoff to thus provide optimized engine operation over a range of flight conditions with respect to performance and other operational parameters such as noise levels.
The engine static structure 44 generally has sub-structures including a case structure often referred to as the engine backbone. The engine static structure 44 generally includes a fan case 46, an intermediate case (IMC) 48, a high pressure compressor case 50, a combustor case 52A, a high pressure turbine case 52B, a thrust case 52C, a low pressure turbine case 54, and a turbine exhaust case 56 (FIG. 1B). Alternatively, the combustor case 52A, the high pressure turbine case 52B and the thrust case 52C may be combined into a single case. It should be understood that this is an exemplary configuration and any number of cases may be utilized.
The fan section 20 includes a fan rotor 32 with a plurality of circumferentially spaced radially outwardly extending fan blades 34. The fan blades 34 are surrounded by the fan case 46. The core engine case structure is secured to the fan case 46 at the IMC 48 which includes a multiple of circumferentially spaced radially extending struts 40 which radially span the core engine case structure and the fan case 46.
The engine static structure 44 further supports a bearing system upon which the turbines 28, 18, compressors 26, 16 and fan rotor 32 rotate. A #1 fan dual bearing 60 which rotationally supports the fan rotor 32 is axially located generally within the fan case 46. The #1 fan dual bearing 60 is preloaded to react fan thrust forward and aft (in case of surge). A #2 LPC bearing 62 which rotationally supports the low spool 14 is axially located generally within the intermediate case (IMC) 48. The #2 LPC bearing 62 reacts thrust. A #3 fan dual bearing 64 which rotationally supports the high spool 24 and also reacts thrust. The #3 fan bearing 64 is also axially located generally within the IMC 48 just forward of the high pressure compressor case 50. A #4 bearing 66 which rotationally supports a rear segment of the low spool 14 reacts only radial loads. The #4 bearing 66 is axially located generally within the thrust case 52C in an aft section thereof. A #5 bearing 68 rotationally supports the rear segment of the low spool 14 and reacts only radial loads. The #5 bearing 68 is axially located generally within the thrust case 52C just aft of the #4 bearing 66. It should be understood that this is an exemplary configuration and any number of bearings may be utilized.
The #4 bearing 66 and the #5 bearing 68 are supported within a mid-turbine frame (MTF) 70 to straddle radially extending structural struts 72 which are preloaded in tension (FIGS. 1C-1D). The MTF 70 provides aft structural support within the thrust case 52C for the #4 bearing 66 and the #5 bearing 68 which rotatably support the spools 14, 24.
A dual rotor engine such as that disclosed in the illustrated embodiment typically includes a forward frame and a rear frame that support the main rotor bearings. The intermediate case (IMC) 48 also includes the radially extending struts 40 which are generally radially aligned with the #2 LPC bearing 62 (FIG. 1B). It should be understood that various engines with various case and frame structures will benefit from the present invention.
The turbofan gas turbine engine 10 is mounted to aircraft structure such as an aircraft wing through a mount system 80 attachable by the pylon 12. The mount system 80 includes a forward mount 82 and an aft mount 84 (FIG. 2A). The forward mount 82 is secured to the IMC 48 and the aft mount 84 is secured to the MTF 70 at the thrust case 52C. The forward mount 82 and the aft mount 84 are arranged in a plane containing the axis A of the turbofan gas turbine engine 10. This eliminates the thrust links from the intermediate case, which frees up valuable space beneath the core nacelle and minimizes IMC 48 distortion.
Referring to FIGS. 2A-2C, the mount system 80 reacts the engine thrust at the aft end of the engine 10. The term “reacts” as utilized in this disclosure is defined as absorbing a load and dissipating the load to another location of the gas turbine engine 10.
The forward mount 82 supports vertical loads and side loads. The forward mount 82 in one non-limiting embodiment includes a shackle arrangement which mounts to the IMC 48 at two points 86A, 86B. The forward mount 82 is generally a plate-like member which is oriented transverse to the plane which contains engine axis A. Fasteners are oriented through the forward mount 82 to engage the intermediate case (IMC) 48 generally parallel to the engine axis A. In this illustrated non-limiting embodiment, the forward mount 82 is secured to the IMC 48. In another non-limiting embodiment, the forward mount 82 is secured to a portion of the core engine, such as the high-pressure compressor case 50 of the gas turbine engine 10 (see FIG. 3 ). One of ordinary skill in the art having the benefit of this disclosure would be able to select an appropriate mounting location for the forward mount 82.
Referring to FIG. 4A, the aft mount 84 generally includes a first A-arm 88A, a second A-arm 88B, a rear mount platform 90, a whiffle tree assembly 92 and a drag link 94. The rear mount platform 90 is attached directly to aircraft structure such as the pylon 12. The first A-arm 88A and the second A-arm 88B mount between the thrust case 52C at case bosses 96 which interact with the MTF 70 (FIGS. 4B-4C), the rear mount platform 90 and the whiffle tree assembly 92. It should be understood that the first A-arm 88A and the second A-arm 88B may alternatively mount to other areas of the engine 10 such as the high pressure turbine case or other cases. It should also be understood that other frame arrangements may alternatively be used with any engine case arrangement.
Referring to FIG. 4D, the first A-arm 88A and the second A-arm 88B are rigid generally triangular arrangements, each having a first link arm 89 a, a second link arm 89 b and a third link arm 89 c. The first link arm 89 a is between the case boss 96 and the rear mount platform 90. The second link arm 89 b is between the case bosses 96 and the whiffle tree assembly 92. The third link arm 89 c is between the whiffle tree assembly 92 rear mount platform 90. The first A-arm 88A and the second A-arm 88B primarily support the vertical weight load of the engine 10 and transmit thrust loads from the engine to the rear mount platform 90.
The first A-arm 88A and the second A-arm 88B of the aft mount 84 force the resultant thrust vector at the engine casing to be reacted along the engine axis A which minimizes tip clearance losses due to engine loading at the aft mount 84. This minimizes blade tip clearance requirements and thereby improves engine performance.
The whiffle tree assembly 92 includes a whiffle link 98 which supports a central ball joint 100, a first sliding ball joint 102A and a second sliding ball joint 102B (FIG. 4E). It should be understood that various bushings, vibration isolators and such like may additionally be utilized herewith. The central ball joint 100 is attached directly to aircraft structure such as the pylon 12. The first sliding ball joint 102A is attached to the first A-arm 88A and the second sliding ball joint 102B is mounted to the first A-arm 88A. The first and second sliding ball joint 102A, 102B permit sliding movement of the first and second A-arm 88A, 88B (illustrated by arrow S in FIGS. 5A and 5B) to assure that only a vertical load is reacted by the whiffle tree assembly 92. That is, the whiffle tree assembly 92 allows all engine thrust loads to be equalized transmitted to the engine pylon 12 through the rear mount platform 90 by the sliding movement and equalize the thrust load that results from the dual thrust link configuration. The whiffle link 98 operates as an equalizing link for vertical loads due to the first sliding ball joint 102A and the second sliding ball joint 102B. As the whiffle link 98 rotates about the central ball joint 100 thrust forces are equalized in the axial direction. The whiffle tree assembly 92 experiences loading only due to vertical loads, and is thus less susceptible to failure than conventional thrust-loaded designs.
The drag link 94 includes a ball joint 104A mounted to the thrust case 52C and ball joint 104B mounted to the rear mount platform 90 (FIGS. 4B-4C). The drag link 94 operates to react torque.
The aft mount 84 transmits engine loads directly to the thrust case 52C and the MTF 70. Thrust, vertical, side, and torque loads are transmitted directly from the MTF 70 which reduces the number of structural members as compared to current in-practice designs.
The mount system 80 is compact, and occupies space within the core nacelle volume as compared to turbine exhaust case-mounted configurations, which occupy space outside of the core nacelle which may require additional or relatively larger aerodynamic fairings and increase aerodynamic drag and fuel consumption. The mount system 80 eliminates the heretofore required thrust links from the IMC, which frees up valuable space adjacent the IMC 48 and the high pressure compressor case 50 within the core nacelle C.
It should be understood that relative positional terms such as “forward,” “aft,” “upper,” “lower,” “above,” “below,” and the like are with reference to the normal operational attitude of the vehicle and should not be considered otherwise limiting.
FIG. 6 shows an embodiment 200, wherein there is a fan drive turbine 208 driving a shaft 206 to in turn drive a fan rotor 202. A gear reduction 204 may be positioned between the fan drive turbine 208 and the fan rotor 202. This gear reduction 204 may be structured and operate like the gear reduction disclosed above. A compressor rotor 210 is driven by an intermediate pressure turbine 212, and a second stage compressor rotor 214 is driven by a turbine rotor 216. A combustion section 218 is positioned intermediate the compressor rotor 214 and the turbine section 216.
FIG. 7 shows yet another embodiment 300 wherein a fan rotor 302 and a first stage compressor 304 rotate at a common speed. The gear reduction 306 (which may be structured as disclosed above) is intermediate the compressor rotor 304 and a shaft 308 which is driven by a low pressure turbine section.
The foregoing description is exemplary rather than defined by the limitations within. Many modifications and variations of the present invention are possible in light of the above teachings. The disclosed embodiments of this invention have been disclosed, however, one of ordinary skill in the art would recognize that certain modifications would come within the scope of this invention. It is, therefore, to be understood that within the scope of the appended claims, the invention may be practiced otherwise than as specifically described. For that reason the following claims should be studied to determine the true scope and content of this invention.

Claims

What is claimed is:

1. A gas turbine engine comprising:

a propulsor section including a rotor and a plurality of blades;

an epicyclic gear train having a gear reduction ratio of greater than 2.5:1;

a compressor section including a first compressor and a second compressor axially aft of the first compressor relative to an engine axis, the first compressor including three stages, and the second compressor including a greater number of stages than the first compressor;

a turbine section including a first turbine and a second turbine, the first turbine including two stages, the second turbine including a greater number of stages than the first turbine but including a lesser number of stages than the second compressor, and the second turbine having a pressure ratio greater than 5:1 and driving the rotor through the gear train; and

a static structure comprising a plurality of cases distributed along the engine axis, wherein the plurality of cases includes a first case and a second case positioned aft of the first case relative to the engine axis, the first case includes a first engine mount location, the second case includes a second engine mount location that reacts an engine thrust in operation, wherein the first engine mount location is axially near the gear train relative to the engine axis, and each of the first engine mount location and second engine mount location supporting an engine mount when the engine is mounted; and

wherein the static structure supports a bearing system, and wherein the first turbine, the second turbine and the rotor are rotatably supported by the bearing system.

2. The gas turbine engine as recited in claim 1, wherein the epicyclic gear train is a planetary gear system.

3. The gas turbine engine as recited in claim 2, wherein the second engine mount location is configured so as to not be connected to the first case by any thrust links when the gas turbine engine is mounted.

4. The gas turbine engine as recited in claim 2, wherein the plurality of blades have a fixed stagger angle.

5. The gas turbine engine as recited in claim 2, wherein the propulsor delivers airflow to an entrance of a core flow path, the core flow path extends through the compressor section, and the epicyclic gear train extends axially aft of the entrance relative to the engine axis.

6. The gas turbine engine as recited in claim 5, wherein the second turbine includes no more than six stages.

7. The gas turbine engine as recited in claim 2, wherein the first case is a compressor case.

8. The gas turbine engine as recited in claim 2, wherein:

the second case is a thrust case; and

the turbine section includes a mid-turbine frame axially between first and second turbines relative to the engine axis, and the mid-turbine frame supports bearings of the bearing system within the thrust case.

9. The gas turbine engine as recited in claim 2, wherein the propulsor is a fan, and further comprising:

a first nacelle which at least partially surrounds a second nacelle and the rotor, the first nacelle having a first exit, the second nacelle having a second exit at an axial location aft of the first exit along the engine axis, and the fan communicates airflow into the first nacelle and the second nacelle; and

a bypass ratio of greater than 10:1.

10. The gas turbine engine as recited in claim 9, further comprising a fan pressure ratio of less than 1.45 across the blade alone at cruise at 0.8 Mach and 35,000 feet.

11. The gas turbine engine as recited in claim 10, wherein the second turbine includes no more than six stages, and the second compressor includes seven stages.

12. A gas turbine engine comprising:

a propulsor section including a rotor with a plurality of blades;

an epicyclic gear train having a gear reduction ratio of greater than 2.5:1;

a first spool including a first compressor and a first turbine, wherein the first compressor includes a plurality of stages, the first turbine includes three stages, and the first turbine drives the rotor through the gear train;

a second spool including a second compressor and a second turbine, wherein the second turbine includes two stages, the first turbine includes a greater number of stages than the second turbine but includes a lesser number of stages than the second compressor, and the first and second spools are rotatable about an engine axis; and

a static structure comprising a first case and a first static structure component located forward of a second static structure component, the first static structure component having a first engine mount location, the second static structure component having a second engine mount location, each of the first engine mount location and the second engine mount location supporting an engine mount when the engine is mounted, and the first engine mount location is axially near the gear train relative to the engine axis.

13. The gas turbine engine as recited in claim 12, wherein:

the engine mount includes a forward mount associated with the first engine mount location and an aft mount associated with the second engine mount location;

the aft mount reacts a thrust vector of a thrust load along the engine axis at the second case in operation; and

the forward mount reacts a vertical load and a side load in operation.

14. The gas turbine engine as recited in claim 13, wherein the first case is an intermediate case that at least partially surrounds the gear train.

15. The gas turbine engine as recited in claim 13, wherein the first case is a compressor case.

16. The gas turbine engine as recited in claim 13, wherein the second case is a turbine exhaust case.

17. The gas turbine engine as recited in claim 13, wherein:

the second case is a thrust case; and

the turbine section includes a mid-turbine frame axially between first and second turbines relative to the engine axis, the mid-turbine frame supports a pair of bearings of the bearing system within the thrust case, and the pair of bearings rotatably support the first and second spools.

18. The gas turbine engine as recited in claim 17, wherein the first compressor includes three stages, the first turbine includes no more than six stages, and the second compressor includes seven stages.

19. The gas turbine engine as recited in claim 17, wherein the propulsor section is a fan section including a fan case, the propulsor is a fan, and the fan case surrounds the fan.

20. The gas turbine engine as recited in claim 19, further comprising a bypass ratio of greater than 10:1.